Fuel control for aviation gas turbine power plants



y 950 R. A. NEAL ETAL FUEL CONTROL FOR AVIATION GAS TURBINE POWER PLANTS Filed March 2, 1948 mvwona- EIV EAL CARI. L sfiDLER,JR.

h 4am ATTORNEY Patented May 9, 1950 I FUEL CONTROL FOR AVIATION GAS TURBINE POWER PLANTS Robert Alden Neal, Media, Pa., and Carl L. Sadler, Jr., Rockford, 111., house Electric Corporation, East Pittsburgh, Pa., a corporation of Pennsylvania Application March 2, 1948, Serial No. 12,614

(01. Gil-35.6)

6 Claims.

This invention relates to combustion apparatus, more particularly to fuel supply systems therefor, and has for an object to provide improved apparatus of this character.

Some aviation jet propulsion power plants are equipped with auxiliary combustion apparatus, disposed downstream of the turbine which drives the compressor, and supplied with gases compressed by the compressor and heated by primary or main combustion apparatus disposed upstream thereof. Such auxiliary combustion apparatus may take the'form of a tail burner or after burner and serves to increase the temperature of the exhaust gases prior to their discharge through the exhaust nozzle of the power plant in the form of a propulsion jet. I

The present invention provides a fuel supply system in which a main fuel pump supplies fuel to the main combustion chamber only, and an auxiliary dual fuel pump normally supplies its entire output to the tail burner. The system includes an automatic control valve which func tions upon failure of the main fuel pump to direct the output of one element of the dual auxiliary pump to the main burner so that it may continue to operate, the'output of the other element of the dual auxiliary pump serving to maintain operation of the auxiliary or tall burner, although at a reduced rate.

Therefore, it is an object of the invention to provide separate fuel pumps for a pair of fuelutilizing devices, together with means for dividing the output of one of said pumps between the pair of fuel-utilizing devices upon failure of the other pump.

Another object of the invention is to provide, in a fuel system for a pair of fuel-utilizing devices each having a separate fuel pump, a control valve for directing a portion of the output of one of the fuel pumps to the fuel-utilizing device normally supplied by the other fuel pump, upon fail-,.

ure of said other fuel pump.

These and other objects are effected by the invention as will be apparent from the following description and claims taken in connection with the accompanying drawing, forming a part of this application, in which:

Fig. 1 is a side elevational view, partially in section, of a gas turbine power plant incorporating the present invention, with the fuel syste shown schematically; and

Fig. 2 is an enlarged sectional view of the control valve of F18. 1.

The present invention, although not limited thereto, is particularly adapted for use with a gas assignors to Westing- 2 turbine power plant of the type employed on aircraft'to drive the propeller or an electric generator, or to supply motive fluid for jet propulsion of the aircraft.

Such a plant preferably comprises a streamlined tubular casing having mounted axially therein a compressor adjacent the forward or inlet end, a turbine downstream thereof, and combustion apparatus located between the compressor and the turbine for heating the compressed air and which discharges the hot gases at suitable temperature and pressure to the turbine. The gases on leaving the turbine are discharged through a nozzle provided at the rear of the casing and may aid in propelling the aircraft.

Referring now to the drawing more in detail, the power plant shown in Fig. 1, and indicated in its entirety by the reference character i0, is adapted to be mounted in or on the fulselage or wing of an aircraft with the inlet H or left end.

as viewed in this figure, pointed in the direction of flight.

The plant comprises an outer shell or casing structure l2l2a providing an annular air duct or passage l3 extending fore and aft with respect to the aircraft. This casing has mounted there in, along its longitudinal axis, a fairing cone H adapted to house gearing connecting through a hollow guide vane IS with auxiliaries (not shown), an axial flow compressor I'I, main combustion apparatus generally indicated l8, a turbine l9 which drives the compressor l'l, an auxiliary combustion chamber 20, and a nozzle 2| defined by the rear end of the casing l2a.

Air enters at the inlet II and flows substantially straight through the plant, passing through the compressor H, where its pressure is raised, and into the main combustion apparatus I8, where it is heated. The hot gases, comprising the products of combustion and excess air heated by the combustion, on leaving the main combustion apparatus are directed by suitable guide vanes or nozzles 23 against the blades 24 of the turbine discs 25, then are further heated in the auxiliary combustion chamber 20, and finally are discharged through the propulsion nozzle 2| to propel the aircraft.

By reference to Fig. 1 it will be noted that the compressor and turbine rotors are interconnected by means of a shaft 26 supported by suitable bearings 21 and enclosed by an inner wall structure, generally indicated 28, which protects the shaft and bearings from high temperatures and also defines a portion of the annular air flow amen passage l3 in which the main combustion apparatus I3 is disposed.

The present invention is not limited to the specific details or arrangements of structure thus far described, but is primarily concerned with the fuel supply system for the main and auxil iary combustion chambers l8 and 20, respecpower plant, with the result that failure of. either pump does not interr'uptoperation of the other.

j The main pump 34 directs its output through conduit 4| of the governor 42, where the desired quantity of fuel is directed to the main combustion apparatus 18 through the conduit 43, ex-- ,cess fuel being by-passed to the inlet side of the pump through the return conduit 44.

The dual auxiliary pump 35-35 directs its output through conduits 45 and 46 to the inlet ports 41 and 48, respectively, of the control'valve the main combustion chamber I l, with the result that the, latter is enabled to continue operation, although at a decreased rating.

While the invention has been shown in but one form, it will be obvious to those skilled in the art that it is not so limited, but is susceptible of various changes and modifications without departing from the spirit thereof.

What is claimed is:

1. In an aviation gas turbine power plant hav- 7 ing an air compressor, a gas turbine for driving said compressor, combustion apparatus for heating air compressed by the compressor to serve as motive fluid for the turbine, and an after 50. This control valve 50 comprises a housing 5! having a cylindrical bore or chamber 52 for the valve member 53, the chamber being closed at one end by an integral wall 54 and at the opposite end by a cover 55.

The valve member 53 includes a pair of lands 53 and 5'! separated by an annular groove 58, and is movable longitudinally in its chamber 52 I between a normal position (shown in Fig. 2)

where the valve groove 58 provides communication between the inlet ports 41 and 48 and an outlet port 59 which connects through conduit 1 6l 6ia with the auxiliary combustion chamber or tail burner 20, and an emergency position- (with the valve member against the chamber end wall 54) where land 51 separates the fiow of fuel from the inlet ports, with the fuel from port 48 1 flowing through conduit 6| to the auxiliary combustion chamber and the fuel fromthe port 41 I passing through conduit to the governor 42 l and thence through conduit 43 to the main combustion chamber. to the auxiliary combustion chamber is determined by control 62 which by-passes unwanted fuel through conduit 63 to the inlet side of the The quantity of fuel admitted D l S'-.

A conduit 65 provides communication between 3 the conduit 4|, at the outlet side of the main fuel pump 34, and the interior of the valve chamber 52, below the valve. member 53, thereby sub- 3 jecting the latter space to the discharge pressure of pump 34. Under normal conditions of operation, this pressure on the lower end of the valve member 53 is sufiicient to maintain the valve member in its normal raised position with the spring 61 above the valve member compressed. With thevalve member in this normal position all of the output of the dual auxiliary pump 135-46 is directed to the auxiliary combustion chamber. Upon failure of the main pump 34, the pressure beneath the valvemember 53 will I be reduced and the spring 51 will move the valve ,member to its emergency position where it contacts the chamber end wall 54. member in this emergency position, the output of the dual auxiliary pump 353'6 is divided between the auxiliary combustion chamber 20 and With the valve source to the main combustion chamber, an auxportion of the output of said auxiliary fuel pump from said second fuel-admission means to said first fuel-admission means uponfailure of the main fuel pump.

2. Structure as specified in claim 1, wherein said auxiliary fuel pump is of the dual type with both parts normally supplying the after burner, and the means for diverting a portion of the output of the auxiliary pump comprises valve structure including a. housing defining a valve chamber having a pair of ports for-admission of fuel thereto from the two parts of said dual pump and first and second outlet ports for discharge of fuel therefrom, a valve movable in said chamber between a normal first position where it directs fuel from both admission ports to said first outlet port for flow to the second fuel-admission means, and

to the first fuel admission means.

3. Structure as specified in claim 2, including means responsive to normal operation of the .main fuel pump for providing a first force for maintaining the valve in its normal first position,

and means providing a second force on the valve in opposition to said first force for moving the valve from its normal first position to its emer gency second position upon loss of said first force resulting from failure of the main fuel pump.

4. In a fuel system for an aviation gas turbine power plant of the type having a main combustion chamber upstream of the turbine, considered in the direction of flow of gases through the plant, and an auxiliary combustion chamber downstream of the turbine; a source of fuel, a main fuel pump for supplying fuel from said iliary fuel pump for supplying fuel from said source to the auxiliary combustion chamber, and control means operable automatically, upon failure of the main fuel pump, to direct to the main combustion chamber a portion of the fuel supplied by the auxiliary pump. 5. In a fuel system for an aviation power plant of the type having a main combustion chamber and an auxiliary combustion chamber; a source of fuel, a main fuel pump for supplying fuel from said source to the main combustion chamber at superatmospheric pressure, an auxiliary fuel pump for supplying fuel from said source to the auxiliary combustion chamber at superatmos- SJOQBU.

pheric pressure, and control means responsive to failure of the main fuel pump for by-paseing to the main combustion chamber a portion of the fuel supplied by the auxiliary fuel pump.

0. Structure asspecifled in claim 5, wherein the control means comprises valve structure including a housing defining a valve chamber having a pair of supply Ports for admission of fuel thereto from the auxiliary fuel pump, and first and second discharge ports for discharge of fuel thereupon, valve means movable in said chamber between a first position where it directs fuel from 6 both supply ports to said first discharge port for flow to the auxiliary combustion chamber, and a second position where it directs fuel from one of said. pair of fuel supply ports to said first discharge port for flow to the auxiliary combustion chamber and directs fuel from the other of said pair of fuel supply ports to the second discharge 'port for tlow'to the main combustion chamber. ROBERT ALDEN NEAL. CARL L. SADLER. JR,

No references cited. 

